Gas turbine engine bifurcation located fan variable area nozzle

ABSTRACT

A turbofan gas turbine engine includes a core engine within a core nacelle, a fan nacelle at least partially surrounding the core nacelle to define a bypass flow path and a variable fan nozzle exit area for bypass flow, and a pylon variable area flow system which operates to effect the bypass flow. A method of operating a turbofan gas turbine engine is also disclosed.

REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.15/889,304, filed Feb. 6, 2018, which a continuation of U.S. patentapplication Ser. No. 13/343,964, filed Jan. 5, 2012, which is acontinuation in part of U.S. patent application Ser. No. 12/441,546,filed Mar. 17, 2009.

BACKGROUND OF THE INVENTION

The present invention relates to a gas turbine engine, and moreparticularly to a turbofan engine having a bifurcation which effectivelyvaries a fan nozzle exit area by adjusting a variable area flow systemwithin the bifurcation to selectively vary the bypass area through whichbypass flow may pass.

Conventional gas turbine engines include a fan section and a core enginewith the fan section having a larger diameter than that of the coreengine. The fan section and the core engine are disposed in series alonga longitudinal axis and are enclosed in a nacelle. An annular stream ofprimary airflow passes through a radially inner portion of the fansection and through the core engine to generate primary thrust.

Combustion gases are discharged from the core engine through a primaryairflow path and are exhausted through a core exhaust nozzle. An annularfan flow path, disposed radially outwardly of the primary airflow path,passes through a radial outer portion between a fan nacelle and a corenacelle and is discharged through an annular fan exhaust nozzle definedat least partially by the fan nacelle and the core nacelle to generatefan thrust. A majority of propulsion thrust is provided by thepressurized fan air discharged through the fan exhaust nozzle, theremaining thrust provided from the combustion gases discharged throughthe core exhaust nozzle.

The fan nozzles of conventional gas turbine engines have a fixedgeometry. The fixed geometry fan nozzles are a compromise suitable fortake-off and landing conditions as well as for cruise conditions. Somegas turbine engines have implemented fan variable area nozzles. The fanvariable area nozzle provide a smaller fan exit nozzle diameter duringcruise conditions and a larger fan exit nozzle diameter during take-offand landing conditions. Existing fan variable area nozzles typicallyutilize relatively complex mechanisms that increase overall engineweight to the extent that the increased fuel efficiency typicallyassociated with the use of a fan variable area nozzle may be negated.

SUMMARY OF THE INVENTION

A gas turbine engine according to an exemplary aspect of the presentdisclosure may include a core engine defined about an axis, a gearsystem driven by the core engine, the gear system defines a gearreduction ratio of greater than or equal to about 2.3, a fan driven bythe gear system about the axis to generate a bypass flow, and a variablearea flow system which operates to effect the bypass flow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may include an annularfan variable area nozzle (FVAN).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gas turbine engine may include a gear systemdriven by the core engine to drive the fan. The gear system may define agear reduction ratio of greater than or equal to about 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gas turbine engine may include a gear systemdriven by the core engine to drive the fan. The gear system may define agear reduction ratio of greater than or equal to 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the core engine may include a low pressure turbinewhich defines a pressure ratio that is greater than about five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the core engine may include a low pressure turbinewhich defines a pressure ratio that is greater than five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow may define a bypass ratio greaterthan about six (6).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow may define a bypass ratio greaterthan about ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow may define a bypass ratio greaterthan ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may operate to changea pressure ratio of the bypass flow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may operate to vary anarea of a fan nozzle exit area for the bypass flow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan may be defined for a predefined flightcondition. Additionally or alternatively, the predefined flightcondition may be about 0.8 MACH and about 35,000 feet. Additionally oralternatively, the predefined flight condition may be 0.8 MACH and35,000 feet.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan may include fan blades designed at aparticular fixed stagger angle related to the flight condition.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may operate to adjustthe bypass flow such that an angle of attack of the fan blades aremaintained close to a design incidence at flight conditions other thanthe predefined flight condition.

A gas turbine engine according to another exemplary aspect of thepresent disclosure may include a core engine defined about an axis. Thecore engine may include a low pressure turbine which defines a pressureratio that is greater than about five (5), a fan driven by the coreengine about the axis to generate a bypass flow, and a variable areaflow system which operates to effect the bypass flow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may include an annularfan variable area nozzle (FVAN).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the core engine may include a low pressure turbinewhich defines a pressure ratio that is greater than five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gas turbine engine may include a gear systemdriven by the core engine to drive the fan. The gear system may define agear reduction ratio of greater than or equal to about 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow may define a bypass ratio greaterthan about six (6). Additionally or alternatively, the bypass flow maydefine a bypass ratio greater than about ten (10). Additionally oralternative, the bypass flow may define a bypass ratio greater than ten(10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may operate to changea pressure ratio of the bypass flow. Additionally or alternatively, thevariable area flow system may operate to vary an area of a fan nozzleexit area for the bypass flow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan may be defined for a predefined flightcondition. Additionally or alternatively, the flight condition may beabout 0.8 MACH and about 35,000 feet.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan may include fan blades designed at aparticular fixed stagger angle related to the predefined flightcondition.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the variable area flow system may operate to adjustthe bypass flow such that an angle of attack of the fan blades aremaintained close to a design incidence at flight conditions other thanthe predefined flight condition.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1 is a general schematic partial fragmentary view of an exemplarygas turbine engine embodiment for use with the present invention; and

FIG. 2 is a sectional view through an engine pylon of the engine of FIG.1 at line 2-2 to illustrate a variable area flow system.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates a general partial fragmentary schematic view of a gasturbofan engine 10 suspended from an engine pylon P within an enginenacelle assembly N as is typical of an aircraft designed for subsonicoperation.

The turbofan engine 10 includes a core engine within a core nacelle 12that houses a low spool 14 and high spool 24. The low spool 14 includesa low pressure compressor 16 and low pressure turbine 18. The low spool14 drives a fan section 20 connected to the low spool 14 through a geartrain 22. The high spool 24 includes a high pressure compressor 26 andhigh pressure turbine 28. A combustor 30 is arranged between the highpressure compressor 26 and high pressure turbine 28. The low and highspools 14, 24 rotate about an engine axis of rotation A.

The engine 10 is preferably a high-bypass geared turbofan aircraftengine. In one disclosed, non-limiting embodiment, the engine 10 bypassratio is greater than about six (6) to ten (10), the gear train 22 is anepicyclic gear train such as a planetary gear system or other gearsystem with a gear reduction ratio of greater than about 2.3 and the lowpressure turbine 18 has a pressure ratio that is greater than about 5.Preferably, the engine 10 bypass ratio is greater than ten (10), the fandiameter is significantly larger than that of the low pressurecompressor 16, and the low pressure turbine 18 has a pressure ratio thatis greater than 5. The gear train 22 is preferably an epicyclic geartrain such as a planetary gear system or other gear system with a gearreduction ratio of greater than 2.5. It should be understood, however,that the above parameters are only exemplary of various preferred gearedturbofan engines and that the present invention is likewise applicableto other gas turbine engines.

Airflow enters a fan nacelle 34 which at least partially surrounds thecore nacelle 12. The fan section 20 communicates airflow into the corenacelle 12 to power the low pressure compressor 16 and the high pressurecompressor 26. Core airflow compressed by the low pressure compressor 16and the high pressure compressor 26 is mixed with the fuel in thecombustor 30 where is ignited, and burned. The resultant high pressurecombustor products are expanded through the high pressure turbine 28 andlow pressure turbine 18. The turbines 28, 18 are rotationally coupled tothe compressors 26, 16 respectively to drive the compressors 26, 16 inresponse to the expansion of the combustor product. The low pressureturbine 18 also drives the fan section 20 through the gear train 22. Acore engine exhaust E exits the core nacelle 12 through a core nozzle 43defined between the core nacelle 12 and a tail cone 32.

The core nacelle 12 is supported within the fan nacelle 34 by a pylonstructure often generically referred to as an upper bifurcation 36U andlower bifurcation 36L, however, other types of pylons and supports atvarious radial locations may likewise be usable with the presentinvention.

A bypass flow path 40 is defined between the core nacelle 12 and the fannacelle 34. The engine 10 generates a high bypass flow arrangement witha bypass ratio in which approximately 80 percent of the airflow enteringthe fan nacelle 34 becomes bypass flow B. The bypass flow B communicatesthrough the generally annular (circumferentially broken only by thebifurcations 36U, 36L) bypass flow path 40 and is discharged from theengine 10 through an annular fan variable area nozzle (FVAN) 42 whichdefines a variable fan nozzle exit area 44 between the fan nacelle 34and the core nacelle 12. The upper bifurcation 36U and the lowerbifurcation 36L, although aerodynamically optimized (best seen in FIG.2), occupies some portion of the volume between the core nacelle 12 andthe fan nacelle 34.

Thrust is a function of density, velocity, and area. One or more ofthese parameters can be manipulated to vary the amount and direction ofthrust provided by the bypass flow B. The upper bifurcation 36Upreferably includes a pylon variable area flow system 50 having apassage 56 defined between a pylon intake 52 and a pylon exhaust 54 toselectively vary the FVAN 42 area through which bypass flow B may pass.Preferably, both the pylon intake 52 and the pylon exhaust 54 arevariable and controlled in response to a controller 58. It should beunderstood that although the upper bifurcation 36U is illustrated in thedisclosed embodiment as having the pylon variable area flow passage 50,the lower bifurcation as well as other pylon structures may likewiseinclude such variable area flow systems.

Referring to FIG. 2, the pylon variable area flow system 50 changes thepressure ratio of the bypass flow B. That is, the nozzle exit area 44 iseffectively varied in area by opening and closing the additional flowarea of the pylon variable area flow system 50 to vary the bypass flowB. It should be understood that various actuators 64, 66 incommunication with the controller 58 may be utilized to operate thepylon intake 52 and the pylon exhaust 54 in response to predeterminedflight conditions. It should be understood that either of the pylonintake 52 and the pylon exhaust 54 may be fixed but it is preferred thatboth are adjustable in response to the controller 58 to control the flowarea through the flow passage 56.

The flow passage 56 is defined around a component duct 55 within theupper bifurcation 36U which provides a communication path for wiringharnesses, fluid flow conduits and other components to the core nacelle12 from, for example, the aircraft wing. It should be understood thatvarious flow passage 56 paths will likewise be usable with the presentinvention.

The pylon intake 52 preferably includes an adjustable intake such as alouver system 60 with empirically-designed turning vanes which mostpreferably have a variation of height to minimize the “shadowing” effectcreated by each upstream louver relative the next downstream louver.

The pylon exhaust 54 preferably includes a variable nozzle 59. Thevariable nozzle 59 may include doors, flaps, sleeves or other movablestructure which control the volume of additional fan bypass flow B+through the FVAN 42.

The pylon variable area flow system 50 changes the physical area throughwhich the bypass flow B may pass. A significant amount of thrust isprovided by the bypass flow B due to the high bypass ratio. The fansection 20 of the engine 10 is preferably designed for a particularflight condition—typically cruise at about 0.8 MACH and about 35,000feet. It should be understood that other arrangements as well asessentially infinite intermediate positions are likewise usable with thepresent invention.

In operation, the pylon variable area flow system 50 communicates withthe controller 58 to effectively vary the area of the fan nozzle exitarea 44 through independent or coordinated operation of the pylon intake52 and the pylon exhaust 54. Other control systems including an enginecontroller, a flight control computer or the like may also be usablewith the present invention. As the fan blades of fan section 20 areefficiently designed at a particular fixed stagger angle for the cruisecondition, the pylon variable area flow system 50 is operated to varythe area of the fan nozzle exit area 44 to adjust fan bypass air flowsuch that the angle of attack or incidence of the fan blades aremaintained close to the design incidence at other flight conditions,such as landing and takeoff as well as to meet other operationalparameters such as noise level. Preferably, the pylon variable area flowsystem 50 is closed to define a nominal cruise position fan nozzle exitarea 44 and is opened for other flight conditions. The pylon variablearea flow system 50 preferably provides an approximately 20% (twentypercent) effective area change in the fan nozzle exit area 44.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The preferredembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

What is claimed is:
 1. A turbofan gas turbine engine comprising: a fanincluding a plurality of fan blades; a gear train; a core engine withina core nacelle, the core engine defined along an engine axis, whereinthe core engine includes a first compressor, a second compressor drivenby a first turbine, and a fan drive turbine that drives the fan throughthe gear train; a fan nacelle at least partially surrounding the corenacelle to define a bypass flow path and a variable fan nozzle exit areafor bypass flow; at least one bifurcation extending between the fannacelle and the core nacelle in a radial direction and extending betweena leading edge and a trailing edge in an axial direction with respect tothe engine axis; a pylon variable area flow system which operates toeffect the bypass flow, wherein the pylon variable area flow systemincludes an adjustable pylon intake, an internal flow passage, and anadjustable pylon exhaust, wherein the adjustable pylon intake includes alouvered system having a plurality of turning vanes adjacent the leadingedge, the adjustable pylon exhaust is adjacent the trailing edge, thepylon exhaust is axially forward of a trailing edge of the fan nacellewith respect to the engine axis, and the internal flow passage isdefined within the at least one bifurcation between the pylon intake andthe pylon exhaust; and a controller coupled to a plurality of actuatorsthat operate the pylon intake and the pylon exhaust in response to thecontroller to cause the pylon variable area flow system to selectivelyvary the variable fan nozzle exit area.
 2. The turbofan gas turbineengine as recited in claim 1, wherein the internal flow passage extendsbetween the pylon intake and the pylon exhaust such that bypass flow inthe bypass flow path enters the pylon intake, then flows through theinternal flow passage, and then exits back into the bypass flow pathdownstream of the pylon intake.
 3. The turbofan gas turbine engine asrecited in claim 2, wherein the at least one bifurcation includes aplurality of bifurcations distributed in the bypass flow path.
 4. Theturbofan gas turbine engine as recited in claim 3, wherein thecontroller operates the pylon intake and the pylon exhaust toselectively vary the fan nozzle exit area in response to a flightcondition.
 5. The turbofan gas turbine engine as recited in claim 4,wherein the gear train is an epicyclic gear train.
 6. The turbofan gasturbine engine as recited in claim 5, wherein each of the firstcompressor, the second compressor, the first turbine and the fan driveturbine includes a plurality of stages, and wherein the fan driveturbine includes a greater number of stages than the first turbine. 7.The turbofan gas turbine engine as recited in claim 6, wherein the fanblades have a fixed stagger angle.
 8. The turbofan gas turbine engine asrecited in claim 6, wherein the pylon variable area flow system operatesto change a pressure ratio of the bypass flow in response to thecontroller.
 9. The turbofan gas turbine engine as recited in claim 8,wherein the first compressor includes a greater number of stages thanthe fan drive turbine.
 10. The turbofan gas turbine engine as recited inclaim 8, wherein the fan drive turbine is a three stage turbine.
 11. Theturbofan gas turbine engine as recited in claim 8, wherein the fan isconfigured for a predefined flight condition, and the fan blades aredesigned at a particular fixed stagger angle related to the predefinedflight condition.
 12. The turbofan gas turbine engine as recited inclaim 11, wherein the predefined flight condition is 0.8 MACH and 35,000feet.
 13. The turbofan gas turbine engine as recited in claim 11,wherein the fan drive turbine drives both the first compressor and aninput of the gear train.
 14. The turbofan gas turbine engine as recitedin claim 13, wherein the pylon intake is established along a sidewall ofthe respective bifurcation at a position that is spaced apart from theleading edge.
 15. The turbofan gas turbine engine as recited in claim14, further comprising an annular fan variable area nozzle located atthe trailing edge of the fan nacelle which defines the fan nozzle exitarea.
 16. The turbofan gas turbine engine as recited in claim 15,wherein the pylon exhaust includes a variable nozzle having a moveableflap.
 17. The turbofan gas turbine engine as recited in claim 16,wherein the predefined flight condition is 0.8 MACH and 35,000 feet. 18.The turbofan gas turbine engine as recited in claim 17, wherein theplurality of turning vanes have a variation of height.
 19. The turbofangas turbine engine as recited in claim 18, wherein the plurality ofturning vanes are pivotable into the bypass flow path.
 20. A method ofoperating a turbofan gas turbine engine comprising the steps of: drivinga fan connected to a first spool through a gear train, wherein the fanincludes a plurality of fan blades, the first spool includes a firstcompressor and a fan drive turbine, a fan nacelle at least partiallysurrounds a core nacelle to define a bypass flow path and a variable fannozzle exit area for bypass flow, the core nacelle houses the firstspool and a second spool, the second spool includes a first turbine thatdrives a second compressor, and at least one bifurcation extends betweenthe fan nacelle and the core nacelle in a radial direction and extendsbetween a leading edge and a trailing edge in an axial direction withrespect to an engine axis to support the core nacelle within the fannacelle; varying a variable area flow system in communication with aninternal flow passage through the at least one bifurcation to change avolume of the bypass flow through the at least one bifurcation andcontrol the bypass flow through the variable fan nozzle exit area inresponse to a flight condition, wherein the variable area flow system ismovable toward an open position to increase bypass flow through thevariable fan nozzle exit area and toward a closed position to decreasebypass flow through the variable fan nozzle exit area; and wherein thevariable area flow system includes a variable pylon intake and avariable pylon exhaust both established along the at least onebifurcation, the internal flow passage extends between the pylon intakeand the pylon exhaust, and the pylon intake includes a plurality ofturning vanes moveable in response to an actuator.
 21. The method asrecited in claim 20, wherein the step of varying the variable area flowsystem includes communicating the bypass flow from the bypass flow pathinto the pylon intake, then communicating the bypass flow through theinternal flow passage, and then ejecting the bypass flow back into thebypass flow path downstream of the pylon intake.
 22. The method asrecited in claim 21, wherein the step of varying the variable area flowsystem includes pivoting the plurality of turning vanes into the bypassflow path to establish the open position.
 23. The method as recited inclaim 21, wherein the at least one bifurcation includes a plurality ofbifurcations.
 24. The method as recited in claim 23, wherein the step ofvarying the variable area flow system includes moving the variable areaflow system toward the closed position in response to a cruise flightcondition.
 25. The method as recited in claim 24, wherein the step ofvarying the variable area flow system includes moving the variable areaflow system toward the open position in response to a non-cruise flightcondition.
 26. The method as recited in claim 25, wherein the fan isconfigured for a predefined flight condition, the fan blades aredesigned at a particular fixed stagger angle related to the predefinedflight condition, and the predefined flight condition is 0.8 MACH and35,000 feet.
 27. The method as recited in claim 26, wherein the step ofvarying the variable area flow system occurs such that an angle ofattack of the fan blades is maintained close to a design incidence atflight conditions other than the predefined flight condition.
 28. Themethod as recited in claim 27, wherein the gear train is an epicyclicgear train, wherein each of the first compressor, the second compressor,the first turbine and the fan drive turbine includes a plurality ofstages, and wherein the fan drive turbine includes a greater number ofstages than the first turbine.
 29. The method as recited in claim 28,wherein the plurality of turning vanes have a variation of height, thepylon exhaust includes a variable nozzle having a moveable flap, and thestep of varying the variable area flow system includes pivoting the flapinto the bypass flow path to establish the open position.
 30. The methodas recited in claim 29, wherein the step of varying the variable areaflow system includes pivoting the plurality of turning vanes into thebypass flow path to establish the open position.